Aircraft automatic pilot



Sept. 5, 1961 G. F. JUDE ET AL AIRCRAFT AUTOMATIC PILOT 6 Sheets-Sheet lFiled May 14. 1959 ATT RNEY Sept. 5, 1961 G. F. JUDE ET AL AIRCRAFTAUTOMATIC PILOT 6 Sheets-Sheet 2 Filed May 14. 1959 Sept. 5, 1961 G. F.JUDE ET A1.

AIRCRAFT AUTOMATIC PILOT Sept. 5, 1961 G. F. JUDE ET AL AIRCRAFTAUTOMATIC PILOT 6 Sheets-Sheet 4 Filed May 14. 1959 TONOTV.

Sept. 5, 1961 G. F. JUDE ET AL AIRCRAFT AUTOMATIC PILOT 6 Sheets-Sheet 5Filed May 14, 1959 Owwllwll We TDNOTV..

Sept. 5, 1961 G. F. JUDE ETAL AIRCRAFT AUTOMATIC PILOT 6 Sheets-Sheet 6Filed May 14, '1959 nited States Patent O 2,998,946 AIRCRAF` AUTOMATICPILOT George F. Jude, Fresh Meadows, and Harry Miller,

Westbury, NY., assignors to Sperry Rand Corporation, a corporation ofDelaware Filed May 14, 1959, Ser. No. 813,097 4 Claims. (Cl. 244-77) Thepresent invention relates generally to improvements in automatic controlsystems for aircraft and more particularly to an automatic pilot foraircraft of the character set forth in co-pending application Serial No.571,813 filed March l5, 1956, for Aircraft Automatic Pilots, and thepresent application constitutes a continuation-inpart of thatapplication. The present application contains a more detailed disclosureof the automatic pilot of the referenced parent application and isdirected to improvements not specifically disclosed therein.

rthe various objects and features of the `improved automatic pilot ofthe present invention will become apparent from the followingdescription of the preferred embodiment thereof taken in connection withthe accompanying drawings, wherein,

Fl'GS. 1A, 1B, and 1C together illustrate the present automatic pilot inblock diagram form; and

FIGS. 2 through 14 are schematic wiring diagrams of the interlockcircuits employed for switching the automatic pilot from one mode ofoperation to `another in respouse to operation of manually controlledselector switches, maneuver command switches, and the like. For thereaders convenience, the general function of each of the separatelyillustrated .interlock circuits are suggested by a figure title.

The following specification will be divided into sections in accordancewith the Various modes of operation available to and selectable by thehuman pilot. Generaily these modes are Oif or Disengage, Yaw Damper,Autopilot Engage, Manual Maneuver, and Path Control. Sub-modes ofoperation under the latter are further available such as AltitudeControl, Radio Guidance (VOR/LOC), and Approach (Glide Path). For anillustration of the outward appearance of the pilots control whereby theforegoing modes of operation may be selected, reference may lbe made toFIG. 7 of the aboveidentified parent application. During the descriptionof the autopilot in its various modes of operation, the functioning ofthe interlock circuits, which in some cases automatically changes themode of operation of the autopilot and which also prohibits incompatiblesignals from inuencing pilot operation in any particular mode, will beset forth.

ln the drawings, all relays are considered to be deenergized and allrelay-controlled switches are accordingly illustrated in theirdeenergized positions. The switch contacts represented by an arrow ortriangle denote those contacts which the switch armatures engage whentheir controlling relays are energized.

General Referring now to FIGS. 1A, 1B, and 1C, which are so arrangedthat they may be conveniently fixed together, it will be seen that theautopilot comprises generally three channels--a rudder channel 100, anaileron channel 290, and an elevator channel 300. Each of these rudder,aileron, and elevator channels may be broken down generally into threesections, namely, a short period stabilization section designated `101,201, and 301, respectively, a long period stabilization sectiondesignated 102, 2M, and 302, respectively, and a command section, thelatter being further broken down into a manual input subsection 2tl3 forlateral command and 303 for longitudinal command, and an externalreference input subsection designated 409 for the lateral externalcommand and 401 and 401 for longitudinal external commands.

Each of the three channels terminates in a substantially identicalservomotor system 10ft, 264.-, and Stia for controlling the movement ofthe rudder, aileron, and elevator, respectively. These servomotorsyst-ems are the same as those described in detail in theabove-identified parent appication and will not be described in detailherein. Suftice it to say that they are Velocity type servos, that is,during the normal operation of the automatic pilot the servos are causedto drive in a direction and at a rate proportional to the direction andmagnitude of the input signal supplied thereto. This is accomplished bymeans of a rate generator driven by a servomotor whose output is fedback to the input of a high gain servo amplifier in opposition to thecontrol signal applied to the amplifier. It will be noted that there isno position feedback during normal operation. The aircraft loop isclosed by means of direct measures of craft angular acceleration, i.e.,the surface deflection is controlled by the angular acceleration suchdeflection produces on the aircraft. Such arrangement providessubstantially optimum control of the aircraft under all flightconditions. As will be described below, during the Disengage mode ofoperation, servomotor position information is required and for thispurpose a synchro is clutched to the motor output in this mode. Theelevator servo system includes a trim tab control system 304 foroperating the trim tab of the aircraft for automatic trim purposes asalso described in the parent application.

As in the parent application, automatic short period stabilization ofthe aircraft about each of the crafts primary axes while under controlof the automatic pilot is provided by a pair of linear accelerometersspaced a distance apart along each axis and their combined outputs areso connected that a signal is produced which is proportional to theangular acceleration of the line joining the accelerometers about anaxis at right angles to such line. Also, as in the parent application,one of the accelerometers in the yaw and pitch channel is mounted asnear as possible to the CG. of the aircraft and its output isolated insuch a manner from its companion accelerometer that such output is ameasure of the lateral and vertical linear accelerations of theaircraft, respectively. The manner in which the accelerometers aremounted in the aircraft and electrically connected is illustrated indetail in the parent application. In the present application, referencecharacter designates the forward yaw accelerometer and 106 designatesthe C G. lateral accelerometer, the output of the latter being combinedwith the output of the former to provide the yaw angular accelerationsignal, and the output taken `alone constituting the lateralacceleration signal. Similarly, reference characters 365 and 3-36designate respectively the forward pitch accelerc-meter and the C.G.vertical accelerometer respectively, the output of the latter beingcombined with the output of the former to provide the pitch angularacceleration signal, and the output of the latter alone supplying thevertical acceleration signal. ln the roll channel, reference characters205 and 2% designate respectively the upper and lower rollaccelerometers, the combined output of which provides a measure of theangular acceleration of the craft in roll, i.e., about the craftfore-and-aft axis.

The long period displacement stabilization of the craft about itsprimary axes is provided by vertical gyro 6 which supplies roll attitudedata from output 267 and pitch attitude data from output 307. Thisvertical gyro may be a conventional gyro having, for example, liquidlevels for maintaining the spin axis thereof slaved to gravity vertical.The gyro may also be provided with a quick erecting device operable whenthe pilot is initially turned on. Furthermore, it may be provided withsuitable means 3 for cutting out its erection under conditions where theeffect of gravity vertical has been displaced due to craftaccelerations. A schematic representation of a vertical gyro suitablefor use herein is illustrated in the parent application.

Long period displacement stabilization is provided in the presentautomatic pilot by a directional gyroscope N7 which may be suitablyslaved to the magnetic meridian. This gyro may likewise be equipped withquick erection devices and means for cutting off its slaving duringperiods of flight, wherein accelerations may affect the magnetic datasensor to which the gyro is referenced. In the absence of externalcommands, the attitude displacement signals provided by the vertical anddirectional gyros are applied unmodified to the surfaces of all systemsas will become apparent below.

All external commands applied to the automatic pilot are supplied to theservomotor systems through command computers which function generally tosmooth the input command signal so that abrupt transients are eliminatedor in some modes as electromechanical ti e integrators. rthus, commandseffecting movement of the craft in the vertical direction are applied tothe elevator servo system through a pitch comand computer 368, andcommands effecting lateral movement of the craft, i.e.,

turns, are supplied to the aileron surface servomotor system through aroll command computer 29S. A yaw cornmand computer lt is operable tofollow-up on the directional gyro output during turn commands.

Pitch commands to the autopilot may be applied from three sources; apilot-controlled pitch-rate command 309, displacement of which from adetent position commands a rate of change of pitch attitude whereby theresponse of the craft to such command is identical to that which wouldnormally occur by the pilots moving his control column. The secondexternal pitch command is supplied by an altitude control 314%, thedetails of which are illustrated in copending application Serial No.571,788 filed March l5, 1956, for Longitudinal Axis Control System forAircraft, and assigned to the same assignee as the present application.The third source of pitch command signal is the glide slope receiverStil which supplies signals proportional to the displacement of thecraft from the radio-defined glide path to the radio coupler 460 whichconverts the output of the receiver 311 to signals acceptable by theautomatic pilot. It should be noted that the altitude control 3l@ andglide slope control 311 constitute apparatus for defining a ight path inthe Vertical plane which the aircraft is caused to follow.

Turn commands for the automatic pilot may be supplied from threesources. A pilots turn knob 299 which, like the pitch knob Sti?,provides an output which commands a rate of turn proportional to thesense of displacement of the turn knob from its detent position. Again,the craft response to such command is identical to that which wouldoccur upon turning of the pilots wheel. The second turn command sourceis a pilots magnetic heading selector 2M. Through this command the pilotmay select any desired magnetic heading and the craft will smoothly turnto and thereafter maintain the magnetic heading so selected. The thirdsource of turn command is the output of a VOR/LOC radio receiver 2liwhich, like glide path `receiver 311, provides a signal in accordancewith the displacement of the craft from a radio defined course. Theoutput of the lateral receiver 211i is modified, in a manner to be fullydescribed below, in radio coupler 4%, the output of which is applied tothe automatic control system to cause the craft to approach andthereafter maintain that radio beam which is tuned in on the receiver.

While there is no main Off-On switch for the autopilot, it iseffectively turned on by closure of three circuit breakers. Thesecircuit breakers are shown in FIGS. 2 and 5. An instrument circuitbreaker labeled Inst. CB, an autopilot circuit breaker labeled APCBf anda yaw damper circuit breaker labeled-YDCB As illustrated, closure of theinstrument circuit breaker serves to energize the vertical gyro and itsassociated quick erection system as well as to energize other componentsof the autopilot requiring fixed phase A.C. Similarly, closure of theautopilot circuit breaker supplies energization current for the roll andpitch accelerometer pairs as well as other component excitation such asroll and pitch servo fixed phaseV energization, roll and; pitch commandcomputer fixed phase energization, etc. At the same time, the autopilotcircuit breaker controls an undervoltage sensor which, upon a loss ordropping of AC. power below a predetermined minimum, serves todeenergize relay K-lZti which automatically disengages the roll andpitch servo systems from their respective control surfaces. Closure ofthe yaw damper circuit breaker supplies excitation current for the yaWaccelerometer pair as well as other component excitation necessary foryaw damper operation, for example, energization of the rudder servosystem fixed phase supply. The yaw damper circuit breaker throughrectifier R1 supplies D.C. voltages to the autopilot and yaw damperengage interlock circuitsJ the operation of which will be describedsimultaneously with the description of the autopilot in its variousmodes of operation. Of course, the autopilot may be completely disabledby opening all of these three circuit breakers.

Disengage mode With all of the pilot circuit breakers closed but theservos disengaged, all the sensors are operative to detect movement ofthe craft about its primary axes; the yaw, roll, and pitchy commandcomputers lti, 2%, and 393 are all placed in follow-up on the referencesignals provided by displacement gyros, i.e., the directional gyro 107and vertical gyro 6; and the surface servo systems are placed inposition follow-up on any signals applied to their input amplifiers.Thus, at any time during the disengage mode the human pilot may placethe aircraft in the engage mode of operatic-n Without abrupt transients.The disengage mode is automatically instituted by placing the servoengage switch, schematically illustrated at 9 in FiG. 2, in its Offposition. In This position selector switch arms lf2, 212, and 3l?.terminate on grounded contacts an arm it! terminates on an open Contact.Under these conditions tne rudder, aileron, and elevator engagesolenoids RES, AES, and EES are deenergized the clutches controlledthereby are open. it will be noted that the direct link between theengage switches and the engage solenoids establishes an importantfeature in that actuation of the engage switch to its i positionimmediately opens the circuit to the solenoids, thereby immediately andpositively disengaging the servos should an emergency arise.

As stated, thefservos are placed in position foilow-up on any inputsignal to the servo amplifier during the disengage mode. This isaccomplished through rudder, aileron, and elevator engage synchrosolenoids RSS, ASS, and ESS whichremain deenergized since the switch lt?is on an open terminal and the main autopilot engage solenoid K-NZD andyaw damper solenoid K-ZS are both deenergized, thereby-opening switchesl-tiZBt, 2, 3, and K-IZS-Z. Deenergization of these solenoidsestablishes a direct connection between rudder, aileron, and elevatorseiyomotors and position synchros 113, Zi, and 313, respectively. Since,in the disengage mode, pilot engage relay K-ltPZD and yaw damper engagerelay K-lZS are both deenergized, switch K-ILGZD-i is open and relayKlZA is likewise deenergized which, in turn, causes the roll commandcomputer 26S and the pitch command computer 368 to be placed infollow-up on the roll and pitch gyro signals through operation ofswitches K-ltiZA-l and K-IZA-Z. Similarly, the yaw command computer ltis caused to follow-up on the directional gyro` signal through theoperation of yaw clamp relay K-ltl which is normally energized in theengaged modes. Thus, yaw clamp relay switch K-UG-r is closed by virtueof open contact of engage switch K-102B-4 (Fig. 11) energizing the yawfollow-up motor.

With the command computers 108, 208, and 308 in their follow-up mode,the tachometer gradients are suitably changed so that the follow-up israpid. The roll tachometer gradient is changed by virtue of rollfollowup low tachometer gradient deenergized relay K-1S0 (Fig. l2) andits associated switch K-ISO-l (Fig. 1B). Similarly, pitch follow-uptachometer gradient deenergized relay 1919@ (Fig. 7), through itsassociated switch K-19ti-1, places the pitch follow-up 3dS in fastfollowup on the ventical gyro pitch signal by virtue of deenergizedengage relay K-ltlZD The yaw command computer is normally in fastfollow-up. However, it is placed in slow follow-up during coupling toILS beams as will be described hereinafter.

Also, in the disengage Inode, disengage relay switch K-ltlZA- operatesto short out any signal accumulated by tne roll integrator 110,

Yaw dam per in the yaw damper mode of operation of the autopilot of thepresent invention, the aircraft is under the manual control of the pilotand the yaw damper portion of the automatic pilot only provides dampingof yaw motions of the craft. The yaw damper is especially designed todamp out any Dutch Roll tendency of the airplane during manual iiight, acondition which is inherent in most modern aircraft having severelyswept wings. The yaw damper is so designed that the human pilot does notnormally know that the yaw damper is operating. ln pilot commandedturns, however, he may be conscious that the yaw damper is helping himsince he need not apply any rudder pedal to coordinate his commandedturns. All he need do is to turn his aileron control wheel in thedirection it is desired to turn and in an amount dependent upon thedesired turn rate. The manner in which this is accomplished will becomeevident as the description of the yaw damper mode continues.

To institute the yaw damper Inode of operation, the pilot moves hisservo engage switch 9 to the Yaw Damper position (Fig. 2). It will benoted that rectifier 8 supplies two DC. voltages. One supply is fused asby fuse F and the other is not. The fused D.C. energizes relay I-Sttlwhich serves as a DC. voltage sensor and normally holds its associatedswitch K-500-1 in its energized or down position. The non-fused supplyenergizes only that portion of the autopilot and yaw damper engageinterlock circuit which is necessary for yaw dampcr operation. if afailure in the automatic pilot interlock channels blows fuse F, thenonfused supply is available to operate the yaw damper. With switch 9 inits yaw dam er position, fused DC. is applied through the pilots andco-pilots emergency release buttons P and CP to a yaw damper engagerelay K-128 and an engage switch holdinfI solenoid 12. The function ofthe engage switch holding solenoid -12 is to provide a rnechanicallyreleasible lock on the engage switch 9 such that with the switch in theyaw damper position, should the solenoid become deenergized, the engageswitch will snap automatically through spring action to its offposition, thereby placing the autopilot in the disengage mode.lncrgization of yaw damper relay I-128 serves lto close its associatedswitc. lli-1284, which energizes the rudder engage solenoid RES throughswitch arm 112 to thereby engage the rudder servomotor with the aircraftrudder. Simultaneously, the rudder synchro solenoid RSS disengages therudder synchro 113 from the servo system (Fig. 1C) so that the onlyfeedback to the servo amplifier is that from the servo tachometer.Clutching of the rudder engage clutch serves to close rudder microswitchRMS, thus establishing a circuit which will energize the engage switchholding solenoid 12. If for some reason the rudder engage solenoid doesnot function, or should any other malfunction occur which would cause adeciutching of the rudder servomotor from the surface, microswitch RMSopens thereby deenergizing engage switch holding solenoid 12 and causingswitch 9 to drop to its olf position.

The foregoing interlock operation places the rudder servo system underthe control of yaw accelerometers and 166, as well as the rollaccelerometers 2.65 and 296. The two basic sensors for yaw damping arethe yaw accelerometers. Their combined outputs, proportional to yawangular acceleration, are applied to an amplifier demodulator 114, theD.C. output of which is applied in two branch circuits, one containing alead circuit and filter (shown in detail in the pitch channel 301) whichfunctions as a blocking network and noise filter in the accelerationsignal due to power supply fluctuations and/ or to body-bending effects.The other branch leads the acceleration signal to a lag-lead filter 116,the lag portion of which serves to develop a rate component therefrom,while the lead portion thereof serves to again suppress noise and tocompensate for any body-bending effects. The details of these circuitsare shown in the pitch channel. The angular `acceleration and ratelterms are combined and modulated in demodulator 117 and the resultantoutput is applied to the input of the rudder servo amplifier 11h whichis a very high-gain amplifier yas described in the parent application.After amplification therein, the signal is used to control the rudderservomotor 119 which drives the rudder at a rate proportional to themagnitude of the amplifier input by virtue of tachometer feedback fromrate generator 12d. Thus, angular accelerations of the craft about theyaw axis serve to drive the control surfaces at a rate proportionalthereto, thereby producing a yaw acceleration of the craft in oppositionto that produced by the disturbing moment which initially caused theangular acceleration. The rapid response of the surface to angularaccelerations greatly increases the apparent inertia of the airplane,thereby minimizing attitude deviations in turbulent air.

As stated, through the use of additional signals, turns produced by thehuman pilot are coordinated through yaw damper operation. Long term orsteady state coordination is provided by CG. located accelerometer 166.Its output is parameter controlled as at 109, rendered operative in theyaw damper mode by relay switch K-128-3, and is then applied todemodulator 121, the D.C. output of which is smoothed as by lag filter122 and remodulated at 123. The output of the demodulator is appliedthrough a direct connection to the rudder servo system input and also toyaw damper integrator 124 whose output, in turn, is combined with theangular acceleration signal and applied to the servo amplifier 113. Thedirect feed of lateral acceleration to the rudder servo together withthe yaw damper integrator output selves to provide the long term orsteady state coordination during human pilot initiated turns. Short termor transient miscoordination is substantially eliminated by driving therudder as a function of the rate of roll of the aircraft. In otherwords, coordination is provided during turn entry and turn exit or inthe event of roll producing gusts during the turn. The roll rate signalis developed as follows. The Lipper and lower accelerorneters 2415 and206 provide a signal proportional to roll angular acceleration as in theparent application, this signal being amplified and demcdulated at 214,the D.C. output of which is branched to lead circuit and lilter 215 andlaglead filter 216. Lead circuit and filter 215 serves as a :blockingnetwork and noise filter as in the rudder channel, while lag-lead filter216 serves to derive a rate component from the acceleration signal andlikewise to suppress noise as above. The roll rate component isparameter controlled at 216 and combined with the angular accelerationsignal in demodulator 217, the output of which is applied to the inputof rudder servo 1S through the yaw damper integrator 124 after beingparameter adjusted at 217. Through the operation of integrator 124, theroll signal applied to the rudderV servo amplitier varies in accordancewith thel rate of roll of the aircraft. The parameter controlling theroll rate term provides transient coordination under all flightconfigurations.

Since the aileron servo is disengaged, the roll angular accelerationsignal will be ineifective in the aileron channel. It will be noted inregard to the interlocks in the yaw damper mode that no importantprecautions need be taken prior to engaging the rudder to its servosystem inasmuch as it is only the higher derivative craft movements thatare being damped out. ,In other words, the yaw damper mode may beinstituted at any time, no matter what the aircrafts attitude might beat that time.

Autopilot engage Before discussing the autopilot mode of operation ofthe autopilot of the present invention, it should be noted that thosecircuits operative in the yaw damper mode just described operate in anidentical manner in the autopilot mode so that a description thereofwill not be repeated.

In order to place the autopilot in the full autopilot mode, engageswitch 9 is thrown to its Autopilot position, thereby moving switch arml@ to its uppermost Contact, establishing DC. power at this point.However, before the full autopilot may be engaged, certain conditionsmust be satisfied. When all these conditions are satisedand the pilotinterlock loop isV closed, the autopilot engage relay l-IZD isenergized, thereby permitting operation of the automatic pilot in allits possible modes of operation. The conditions which must be fullledare as follows. There must be sutiicient D.C. power as determined byswitch K-i'l-2 controlled from undervoltage relay K-Stlt); the verticalgyro must be fully erected and operating normally as determined byswitch K-StlZ-l controlled from quick erection relay K-302 (FIG. 5).Also, sufficient A.C. voltage must be present in the system asdetermined by A.C. undervoltage sensor '7 and relay K424i normallyenergized thereby. Switch K-lZG-l determines if this condition issatisfied before the pilot can be engaged. A further condition whichmust be fuliilled before pilot engage relay K-lZD will be energized isthat all autopilot automatic cut-olf relays will be balanced, ie., thatthere is no condition in the autopilot which would cause its automaticcut-oif. The details of such automatic cut-olf or safety monitor systemare disclosed in cop-ending application Serial No. 623,592 forMonitoring Device for Automatic Pilot Systems, tiled November 21, 1956in the names of H. Miller and R. H. Parker, and assigned to the same`assignee as the present invention. For simplicity of illustration,these automatic cut-off relays have been illustrated generally by asuitably labeled block i3.

Also, the autopilot cannot be engaged at bank angles greater than 35.Such condition is determinedby bank detector relay .l-lili (FIG. 1B).This relay is energized through the operation of the roll commandcomputer 20S as, during the disengage mode, it follows up on thevertical gyro roll signal supplied by roll synchro 218 on vertical gyro6 and synchro 2M on the roll computer output shaft. An extension of theroll computer shaft positions a switch arm 226 along an arcuate sector221 which subtends an angle of 35 on each side of zero roll referenceposition. D.C. current is applied to the wiper 22h such that if thewiper is off the Contact strip, bank detector relay K-IM is deenergizedand switch K-ll41 (FIG. 2) will remain open, thereby preventingengagement of the pilot. Of course, if the bank angle is less than plusor minus 35, switch K-114-1 is closed permitting pilot engagement,provided the other engaged conditions are fulfilled.

A still further condition Whichmust ibe satisfied before thegautopilotengage relay will become energized is'thatv the manual pitch andturnlcontrollers must be in their no command or detent position. This isaccomplished in the engage interlock circuit for relay .Ii-102D bynormally closed relay switches K-ltM-B-l and K-iilo-l which, in turn,are controlled by turn knob detent relay K-liB and pitch knob detentrelay .l-166, illustrated in FIG. 3. As shown in this iigure, should theturn knob 14 be out of detent, switch l5 will be closed therebyenergizing relay K-itMB and opening engage interlock circuit switchK-B-l. Similarly, should the pitch knob 16 be out of detent, switch i7will be closed and relay K-Itl energized, thus opening K-iiid-l in theengage interlock circuit (FIG. 2).

In the pilot engage interlock circuit of FIG. 2, the engage lrelayswitch K-IZC-l is shown in parallel with the bank angle switch K-lM-land the detent switches K-ltleB-l and K-lil-I. This switch is closedupon enerigization of the main pilot engage relay H4021) and serves toinsure that subsequent operation of the bank angle and detent relayswill not disengage the pilot.

With all of the foregoing conditions satisfied, the pilot engage relayK-102D is energized, which energization controls a number of parallelconnected engage relays K-JLZA, -B, and -C through the closure ofenengage relay switch K-llGZD-l. Energization of K H2B closes gangedswitches K-1ti2B-l, 2, and -3 which, in turn, energize the rudder,aileron, and elevator synchro solenoids RSS, ASS, and ESS to declutchthe surface position synchros M3, 2l3, and 3l3 from the servomotoroutputs and simultaneously, by virtue of switch arms 112, 2i2, and 312being in their uppermost position, energize the rudder, aileron, andelevator engage solenoids RES, AES, and EES, thereby effecting directconnection from the servomotors to the control surfaces. It will benoted that if the craft had previously been in the yaw damper mode, therudder synchro solenoid and rudder engage solenoids will remainenergized through the operation of engage switch K- 10213-3, the yawdamper relay switch K-l2S-2 having been open upon switching from theyaw' damper to the autopilot mode. With the rudder aileron and elevatormicroswitches RMS, AMS, and EMS all closed, a circuit is establishedtoengage switch holding solenoid l2. Thus, should any malfunction occurwhich would cause any of the servo clutches to separate, the aifectedmicroswitch will open, deenergizing engage switch holding solenoid 12and allowing the engage switch 9 to drop to either its yaw damperposition or off position, depending upon which servo experienced thefailure. The engage switch is so designed that, should the malfunctionnot atfect the yaw damper ele-ments, the engage switch willautomatically drop to yaw damper position and remain there. This isaccomplished in part by separate energization of the engage switchholding solenoid l2y through yaw damper contact on switch lh.

Closure of relay K-ltlZA also serves to place roll command computer 208and pitch command computer 308 in their engage positions wherein theywill respond to any input commands thereto. Through the operation ofl-liZA, roll integrator il is also rendered responsive to any inputsignal applied to its input by closure of switch K-162A-3- Thetachometer gradients of the roll command computer 268 and the pitchcommand computer 363 are adjusted for their normal autopilot operationthrough energization of relays K-'t (HG. 12) and K494i (FiG. 7) andclosure of engage relay switches K-lt2B-4 and K-ZD-fi, respectively.

It has been described above that the servos cannot be engaged if theaircraft is in a banked attitude of greater than plus or minus 35. Ifthe autopilot is engaged at bank angles less than 5, the craft willautomatically roll to level flight and the steady state heading, i.e.,the compass heading, at engagement will be maintained. This isaccomplished as follows. Since during the disengage mode yor yaw dampermode the roll computer 268 followed up on the vertical gyro signal, theroll attitude of the aircraft exists in synchro 219 (FIG. 1B) and thesine of this bank angle is present in resolver 222, also positioned bythe roll command computer shaft. Therefore, upon engagement throughswitch K-102A-1, the stored bank signal is applied to the input of theroll command computer and drives the latter to zero the signal, and inso doing causes aileron operation through gyro synchro 213 and rollcommand computer synchro 219 to roll the aircraft to a level attitude.

Referring now to FIGS. 5 and 1B, a pair of liquid level switches 223 areprovided and are responsive to lateral accelerations of the craft suchas would be produced by a bank angle of about 5 These switches arearranged on the gimbal of the roll gyro as represented by the dottedline connection in FIG. 1B and are normally maintained level by gyrorigidity and hence are very sensitive to lateral accelerations. Theseswitches serve to close a circuit from the instrument circuit breakerthrough a suitable transformer and bridge rectifier 224 to ground, thusenergizing erection cut-off relay K-Slf-i, energization of which servesto cut off the erection on the vertical gyro until the bank angle of thecraft is less than 5 roll, i.e., the lateral acceleration sensed by theswitches drops below a predetermined value. Energization of erectioncut-off relay K34 also energizes bank threshold relay K-110 throughaction of switch K-StM-l (FiG. 6). If, however, the bank angle is under5, gyro erection stays on and the bank threshold relay K-llt) remainsdeenergized. Referring now to FIG. 1l, it can be seen that even with thepilot engage relay closed, yaw command computer 168 is clamped throughthe deenergization of bank threshold relay switch K-M-, thereby allowingyaw command computer relay K-l' to be energized, which disconnects theoutput of the yaw follow-up amplified with its follow-up motor. Also, asshown in FG. ll, gust clamp relay K-111 is also energized, therebysupplying a heading reference signal to the rudder servo system throughgust clamp switch K-1111 and to the roll command computer through gustclamp switch K-111-2. In the foregoing manner, if the pilot is engagedat less than 5 bank angle, the craft will roll level and maintain theheading obtaining at engagement.

lf the autopilot should be engaged at greater than 5" bank angle, thecraft will make a coordinated maneuver to level flight and maintain theheading achieved when the bank angle drops below 5 This roll to levelflight is accomplished in the same manner as before, but while theaircraft is at a bank angle greater than 5 the bank threshold detectorK-iilil will be energized due to the energization of erection cut-offrelay K-304 as a result of closure of one or the other of liquid levels223. Energization of the bank threshold detector opens the switch K-ll11and deenergizes yaw computer camp relay K-llfl, placing yaw follow-up108 in follow-up on the signal from the directional gyro 167.Simultaneously, gust clamp relay K-111 is likewise deenergized so thatno heading reference signal is applied to the rudder servo nor to theroll command computer 208. Thus, as the roll command computer 208reduces the bank angle signal in resolver 222 toward zero, thedirectional gyro remains in follow-up until the bank angle drops below 5at which time the yaw command computer `108 is clamped as describedabove.

Upon engaging the flight control system, the pitch attitude existing atthe time of engagement is maintained and the craft may be returned tolevel flight by operating pitch command knob 16 or by selecting thealtitude control mode. 1f the autopilot is engaged at 0 pitch attitudewith the altitude control off and the pitch knob in detent, pitchfollow-up clamp relay K-191 (FIG. 7) is energized and the pitch commandcomputer 36S is clamped. The output of the pitch command computersynchro 317 is zero at this time by virtue of its follow- CII ing up onthe vertical gyro signal during the disengage mode and any deviationsfrom level flight after engagement will cause the craft to be broughtback to level flight due to the direct connection from vertical gyropitch synchro 31S to the elevator servo system '.ift. Also, if theautopilot is engaged in, say, a pitch down attitude, no pitch signalwill be supplied to the elevator servo system, again due to thefollow-up action of pitch command computer during the disengage mode.

The autopilot is now fully engaged and is conditioned for acceptance ofany maneuver command signals generated either by manually insertedcommands through the turn and pitch knobs 14, 16 or through externallyinserted commands, i.e., radio commands or altitude control commands.

Autopz'loi short term stabilization With the automatic pilot engaged,short term attitude stabilization is provided by the yaw, roll, andpitch angular accelerometer signals produced by the yaw, roll, and pitchaccelerometer pairs. The function of the roll and pitch acceierometerpairs in stabilizing the craft against short period disturbances isexactly the same as that previously described with respect to the yawchannel operating the yaw damping mode and a detailed discussion thereofwill not be repeated. As stated, the upper and lower roll accelerometersignals are combined so as to produce a resultant signal proportional tothe angular acceleration of the craft, the signals being modified aspreviously described and applied to the aileron servo system 294 where asurface rate proportional to the measured acceleration is produced tothereby stabilize the craft against short period roll disturbances. Thepitch angular accelerations are suppressed in the same manner, However,as in the yaw channel, the accelerometer lit-5 mounted at the C G. ofthe craft supplies a separate signal proportional to the verticalacceleration thereat. It will be noted in FIG. 1C that the details ofthe lead filter and lag-lead filter are shown in detail, and it shouldbe understood that the corresponding lters in the other accelerometerchannels are substantially identical, the functions thereof having beenpreviously described in connection with the yaw damper mode ofoperation.

Manual maneuver commands Assume now that the craft is flying straightand level and it is desired to manually maneuver the craft about itspitch axis. For pitch maneuver commands by the human pilot, pitchfunction selector knob 18 must be in its Pitch Knob position (FIG. 4).With the pitch function selector knob 18 in this position, the pitchattitude of the craft may be changed by rotation of spring centralizedpitch rate command knob 16. Rotation of this knob out of detent closesswitch 17 and energizes relay K-i (FIG. 3) which, in turn, deenergizespitch follow-up clamp relay K-191 (FIG. 7) by opening of switch K-lit-Sto thereby unclamp pitch command computer 303 (FiG. 13) by closing ofswitch K-191-1. At the same time, a pitch command signal proportional tothe sense and magnitude of the displacement thereof is generated throughpotentiometer 314, the magnitude and sense of this signal being in turnproportional to the desired sense and rate of change of pitch attitudeof the craft. This signal is applied through now closed pitch detentrelay switch K-1G6-2 to the input of pitch followup amplifier 308 whichenergizes pitch follow-up motor 315 so that it drives in the directionand at a rate proportional to the direction and magnitude of the pitchcommand signal by virtue of the speed feedback voltage from generator316. it will be noted that in this mode relay K-19t (FlrG. 7) isenergized and relay switch K-19ti-1 is moved to its high tachometergradient position, thereby decreasing the rate at which the follow-upmotor can operate as compared with its rate in the disengage mode, thatis, to a rate consistent with craft response character- 'iff istics toinput commands. Itl will be further noted that the motor 315 willcontinue to `drive until the pitch command signal is reduced to zero asby manuaily returning knob 16 to detent. Rotation of the pitch commandcom-V puter synchro 3l? with respect to vertical gyro synchro 313 willproduce an error signal which is supplied to the elevatorV servo system304 to thereby produce a pitch rate of the aircraft proportional to sucherror. in order that the elevator deflection producing the commandedrate is not initially opposed by the angular pitch acceleration whichnormally tends to oppose any rotation of the craft about its pitch aXis,the rate of change of the pitch command signal as measured by follow-upgenerator 316 and hence a pitch acceleration term is applied throughswitch K-lZ-S (deenergized inthe manual pitch command mode) and asuitable isolation transformer 319 to the input of amplifier demodulator32d in the angular accelerationV output channel in opposition to theangular acceleration signal, thereby bucking outV any pitch angularacceleration signal which would oppose the pitch command.

When the desired pitch attitude of the craft has been achieved, thepilot releases or re-centers the pitch knob lo, thereby Zeroing thepitch command signal from po-' tentiometer 314 and closing pitch detentswitch i7. Thus, pitch detent relay K-ti is deenergized and K-iienergized to thereby clamp pitch-command computer 3% at the position itthenV had. Thel vertical gyro thereafter stabilizes the aircraft at thenewly establishedV pitch attitude. The craft may be* returned to levelfiight attitude by an opposite sequence of operation of the pitchcommand'lmob i6, or, if desired, by rotating pitch function selectorswitch 18 to the ALT position. The altitude control mode will bedescribed below.

As stated, two types of manual turns may be accomplished with theautopilot of the present invention; by manual turn rate commandsinserted through operation of turn rate controller 2%9 and bypreselected heading commands throughv heading selector Z'lti. Theselection of these modes is accomplished through the pilots turnfunction selector switch i9 to either the Turn Knob or I-IDG SELpositions, in eachl of which interlock circuits are established wherebycommand turns from these sources are supplied to the autopilot.

Y As shown in FIG. 4, the turn knob position of tion selector switch 19is the normal position, the knob being spring centered to this position.With the knob 19 in this position, relay K444i is energized,conditioning the system for additional operation in either of the radiomodes to be hereinafter described. Turn rate commands are institutedthrough turn knob K4, rotation of which out of detent position closesdetent switch i and energizes relays K1G4A and K-ltMB. Simultaneously, avoltage is generated across potentiometer 225 which is proportional tothe magnitude and sense of such displacement, which voltage is in turnproportional to the magnitude and sense craft rate of turn which it isdesired to make. Energization of relay K-lG4A serves to unclamp the yawcommand computer M8 throughthe opening of switch K-lGGA-l (FIG. ll) withthe resulting deencrgization or" yaw computer clamp relay ifa-179 andgust clamp relay K-lll. Thus, as the aircraft turns in response to theturn command signal, the yaw command computer MBS follows up on anysignal from the directional gyro 197. Energization of relay K-lti-fcloses switch K-lME-i (FIG. 5) which serves immediately to cut off theerection controls of the vertical. gyro and any slaving controls of thegyromagnetic compass system of which the directional gyro M37' andheading selector 2id may form a part. Erection cut-off relay K-Sti.-also, through switch K-StM-', energizes bank threshold relay K-llt)(FIG. 6) which, in turn, opens switch K-.iftl-l in parallel with thegust clamp relay switch K-lilin the yaw command computer clamp and gustclamp circuit (FIG`.' ll); Functionrof these switches will be defunc-1:2 scribed hereinbelow. Relay K-ltlfiB also closes switch K-MMB-Z (FIG.1B) which connects the turn command signal to the roll command computer208.

The turn rate commandV signal applied to roll command computer 268produces a rotation of roll computer motor 225 at a speed determined bygenerator 227 and through an angle proportional to the sine of the bankangle required for the rate of turn commanded. Synchro 219 thus biasesor shifts the bank angle reference provided by synchro 21S on thevertical gyro 6 through an angle in accordance with the rate of turncommand. The signal from synchro 219 is applied to the aileronservometer and causes the craft to bank to that bank angle and thereforeto turn at a rate corresponding to that commanded. Of course, shortperiod roli stabilization continues to be supplied by the upper andlower accelerometers 205 and 206. Any initial opposition to the rollcommand by the roll accelerorneters and the inherent lag inservoresponse produced thereby is not objectionable in this channel. As willbecome evident later on, this lag is actually accentuated in other modesof operation. As explained above, the roll acceleration signal isapplied through yaw damper integrator 124 to apply rudder in a sense tooppose any adverse yaw due to the roll rate. As the craft turns inresponse to the bank angle, the yaw command computer 198 follows up onthe directional gyro iti?. However, since gust clamp relay has closedswitch K-iil-l, any heading error signal exceeding the follow-upcapacity of the yaw command computer is applied to the rudder servosystem 10i) to thereby provide heading stabilization during the turn.Short and long term turn coordination is supplied through the C.G.mounted accelerometer Mio operating directly into the rudder servosystem and through electronic integrator 12d into the rudder servosystem, respectively.

In order to prevent any loss in altitude during the turn, due to thedecrease in vertical lift of the wings by banking of the craft, a liftcompensation signal dependent upon bank angle is inserted into the pitchor elevator channel of the autopilot. It can be shown that the precise0rrathernatical value of this compensating signal may be very closelyapproximated by the expression (l-) where 35 is the bank angle. Thus, afurther winding on resolver 222 rotated by the roll command computermotor 226 provides a signal proportional to cos o. This signal isapplied asan inputV to a (l-o) computer 323 where it is combined with afixed constant voltage. The resultant signal is applied through twobranches to the pitch channel of the autopilot through a suitableisolation transformer 324. In one branch the signal is parametercontrolled as at 326 in accordance with the reciprocal of dynamicpressure q since the value required for lift compensation varies withair speed. This lift compensation signal is applied to the input of theelevator servo system 304 and produces an elevator deflection such as totend to maintain the altitude of the craft constant during the turn.

When it is desired to stop the turn, the pilot rotates the turn knobback to its detent position, the craft immediately rolling to levelflight and, through interlock circuits, it will fly to and maintain theheading obtaining when the bank angle has decreased to below 5. Returnof turn knob 14 to its zero position opens detent switch i5 anddeenergizes relay K-104A and KQLB; Kein-@B4 and -B-4 serving,respectively, to sever the connection between the turn command 299 andthe input tothe roll command computer' 268 (FIG. iB) and to place theerection control relay K-Stift under the control of liquid levels 223(FIG. 5). Also, K- ltilA-i establishes a connection from the D.C. supplyto bank threshold relay switch K-llti-l and gust clamp relay switchK-lll-Il (FIG. l'l). Therefore, as soon as the bank angle' drops toV 5,erection relayv K-Si will ecome deenergized due tothe levelling of theliquid levels 223. At the same time, bank threshold relay K-110 will bealso deenergized, again deenergizing yaw command computer clamp ifi-170and gust clamp K-1i11. Thus, the yaw command computer 108 is againclamped and the directional gyro 2107 provides its heading stabilizationsignal to the yaw and roll channels of the autopilot through gust clamprelay switch K-111-2 (FIG. 1B) to maintain the craft in straight andlevel flight at the heading then obtaining.

If when the craft is rolling out of the bank maneuver and the bank angledrops below but a sudden gust should cause the craft momentarily toincrease its bank angle over 5, the switching sequence above describedwould again unclamp the yaw command computer causing it to follow up onany directional gyro signal. Since a gust is normally of generally shortduration, such momentary operation of yaw command computer isundesirable and means have been provided for allowing the yaw commandcomputer to be unclamped only upon commanded turns rather than upon agust produced bank angle. Upon gust clamp relay K-111 becomingenergized, switch K-111e1 (FIG. l1) in parallel with bank thresholdswitch K-110-1 closes, maintaining a direct D.C. path to the gust clampand yaw corn- Inand computer clamp relays, even though bank thresholdrelay K-110 is energized due to a gust detected by liquid levels 223.

The craft is now back in straight and level flight, holding the headingdefined by the directional gyro 107 and an altitude defined by verticalgyro 6. It will be noted in FIG. 4 that the manual turn knob mode takesprecedence over any other mode of operation of the autopilot through theuse of a suitable electromechanical holding latch 227 coupled withswitch 19, the energization of which is controlled in dependence uponwhether or not turn knob 14 is in or out of detent. For example, if knob19 were in any position other than TURN KNOB and the pilot turned turnknob 14, K-104B would become energized, closing switch K-104B-3` in PIG.4 and energizing holding switch solenoid K-201, automatically centeringselector knob 219 through its centralizing spring.

Command turns may likewise be made through the pilots magnetic headingselector 210. The details of such a selector are shown in theabove-identified parent application and particularly in FIG. 5 of thatapplication. Through this type of turn control, the pilot may turn toany selected magnetic heading and the craft will smoothly bank up into acoordinated turn and then roll to level flight at the magnetic headingso selected. The normal procedure for making a heading type turn is toselect the heading it is desired to ily on the heading selector 210 withthe turn function selector knob 19 in its normal turn knob position. Thesignal so generated will wind up on open contacts of switches K-lOI-land -2 (FIG. 1B). Then, when it is desired to make the turn, the pilotmerely turns the turn function selector knob to the HDG SEL position,the latter setting serving to institute the turn of the craft. In thismanner, the desired heading may be preset and the actual turn of thecraft instituted at any desired time thereafter, a feature which may behighly desirable, especially when holding in a stack preparatory tolauding. Alternatively, the turn function selector knob may be set toheading select and the heading then selected on the selector 210, thecraft smoothly following the heading as it is being selected.

As described in the parent application, setting of the selector 210generates a signal in a synchro coupled with the heading indicator whichis proportional to the difference between the actual heading of thecraft and the desired heading so set. Switching of turn selector knob 19to the heading select position energizes K-101 (FIG. 4) and supplies theheading error signal of heading selector 210 to an amplier limiter 228through switch K-101-1 (FIG. 1B). Also, switch K-101-2 closes a circuitbetween heading selector 210 and the roll integrator 11 whose output iscombined with the heading error signal and applied to the input ofamplifier limiter 228. The integrator serves, in the heading selectmode, to prevent the aircraft from turnin7 due to any persistent trimerror which is reflected in a persistent heading error. It will be notedthat the heading error signal from heading selector 210 is first passedthrough a transient integrator 229 for smoothing out any abrupt motionof the heading selector knob. Energization of relay K403i also arms asmooth engage circuit 230 through switch K-101-3, relay K-ll (FIG. 13)and its associated switch K-118-1 (FIG. 1B). This smooth engage circuitfunctions through a capacitor-charge time type circuit to limit thespeed with which the heading error signal is allowed to be applied tothe roll command computer 2tlg. This is especially desirable in theabove-described preferred manner of making a heading select type turnwhen there might otherwise be a switching transient.

Inasmuch as the magnitude of the error signal from heading selector 210may be very large; as for example when, say, a 180 turn is selected, andbecause a bard; angle proportional to heading error is provided to theaileron servo system, the heading command signal is limited in limiter223 to a magnitude such as not to command an excessive bank angle. Asillustrated in FIG. 1B, the limits imposed on the bank angle may bevaried, depending upon the mode of operation of the aircraft. These willbe discussed later in connection with the radio guidance modes. Theremainder of the operation of the autopilot in the heading select modeis exactly the same as that in the turn knob mode, particularly inregards to the yaw command computer clamping and unclamping and gustclamp operation.

Flight palh control ln this section of the present specification will bediscussed the operation of the autopilot under the influence of commandsgenerated in accordance with departure of the craft from preselected orpredetermined flight paths, that is, ilight paths that are defined bysuch references as barometric data and navigational or other radiobeams.

lf it is desired that the autopilot control the aircraft to maintain adesired pressure altitude, function selector switch 1S is turned to theALT position. As shown in FIG. 4, such switching will energize altitudecontrol relay Y-161 which serves to energize altitude control 310. lnpractice, relay K-161 is in the altitude control unit 310 and functionseffectively to clamp a barometric altitude reference member, such as ananeroid bellows, in the position it had when clamped and any error inthe altitude thereafter will produce a signal output from a suitablepick-oft" device proportional to such deviation. Switch 18 alsoenergizes a second altitude control relay K- which, in turn, closesswitch K-lS-i (FIG. 1B) thereby supplying the altitude error signal tothe autopilot command circuits. Relay K-115 also controls switch K-115-2(FIG. 9) which serves to energize vertical path relay K-112. Operationof this relay in turn opens switch K-112f-1 (FIG. 7) to therebydeenergize pitch follow-up clamp K-191 and allow pitch command computer30S to be responsive to any signal applied to its input amplifier.

At this point it should be stated that if the craft is in altitudecontrol and the pilot wishes manually to change altitude, he need merelyrotate the pitch command knob 16. This action will again close pitchdetent relay K-106 which, in addition to putting in the pitch command asdescribed above, also deenergizes a mechanical interlock relay K-200which automatically returns the function selector switch knob 1S to itspitch knob position. Conversely, if the aircraft is in a climb or divethrough pitch knob operation and it is desired to level out at aparticular altitude, the pilot need merely switch the l functionselector switch lg toits ALTtposition, such switching making theconnections described above to hold the altitude at which the switch lwas operated. The details of the altitude control system are disclosedin the above U.S. application Serial No. 571,788, now Patent No.2,936,134.

As will become evident, when the crait is placed in the altitude controlmode, the signals controlling the elevator servo system include analtitude displacement error signal, a damping term derived from theintegrall of vertical acceleration, and an integral control term derivedthrough integration of any persistent altitude error. in order for suchintegral terms to be derived, certain modifications in the pitch commandcomputer 393 must be accomplished. lt will be remembered that in thepitch maneu ver command mode, the pitch command computer followed uprapidly on the pitch rate command signal. In the altitude control mode(as well as the glide slope mode to be described later), pitch commandcomputer 3% has its gain changed to such an extent that i-t may operateas a long term integrator. Thus, when the altitude mode is selectedthrough selector -switch i8, certain interlocks are elected to makethese changes in the pitch command computer. With selector switch 18 inits altitude position, altitude relay K-llS is energized through Vengagerelay K-lll2C2 (FIG. 4), thereby energizing vertical path relay K-llZthrough engage relay K-lilZD-l (FlG. 9). Energization of K-lllZ opensswitch K-llZ-l' (FIG. 7), thereby deenergizing relay K-ll and unclampingthe pitch command computer 39S. Since the pilot is engaged, i149@ isenergized thereby increasing the amount of rate generator feedbacksignal to the input of the amplifier of the pitch command computer.Energization of K-llZ also closes switch lsI-llZ-Z (FIG. 8') which, inturn, energizes solenoid MG-lill. This solenoid functions to shift gearsin the gear transmission 322. between pitch command computer motor SiSand its output synchro 317. ln the manual pitch rate maneuver mode, ahigh speed gear reduction (for example SOOOzl) is required so that thepitch command computer follows up rapidly on the pitch rate commandsignal. However, to perform as an integrator, the gearing between thepitch command computer motor 315 and output 337 must be a low speedgearing such that considerable rotation of the motor l will produce onlya small rotation of output synchro 35.7 (for example 150001). Thus,operation of the selector knob ll to its ALT position not only switchesin the altitude control 3l@ but also, through the interlocks justdescribed, conditions the pitch command computer 303 to operate as along term integrator of any persistent altitude error.

For the purposes of the present application, the altitude control Sillsupplies an output proportional to the displacement of the craft fromsome reference biarometric altitude, this signal being applied directlythrough an amplifier, lag and limiter circuit 32l directly tothe inputof elevator servo system 3G14. The limiter serves to limit the magnitudeof the pitch attitude commendable by the altitude error signal. Thisdisplacement signal is bucked by a signal from the vertical gyro synchro31S through synchro 317 on the pitch command computer 368 in a suitablesumming circuit. This signal provides short term altitude control which,corrects for gusts and other disturbances. The lag serves to smooth thealtitude error signal. lnertial path dampingis provided in the presentautopilot by a signal proportional to normal accelerations of the craft,i.e., parallel to direction of gravity, as detected by the OG. orvertical accelerometer M35 (FIG. 1C), this signal being applied througha highpass or lead lter to the input of the pitch command computer 368where it is integrated and applied to the pitch servo system as an`altitude rate term. However, with the airplane in a turn, the CG. orvertical accelerometer 1% -would sense the centrifugal accelerationproduced by the turn-and thereforewould tend to noseY thef craft down inorder to reduce such acceleration., Since the vertical `accelerorneterobviously cannot sense the diiference between normal acceleration and avertical acceleration component due to centrifugal force, the otherbranch of the output of the l-cos gb computer is applied' in the outputof vertical accelerometer 106 in such a sense as to cancel only thatcomponent of vertical acceleration produced by turning, i;e., thecentrifugal force component. However, since the l-cos 35 signal is anapproximation, a small ve-rtical acceleration signal may exist ath-ighbank angles. The difference between the norm-al `accelerationl signaland the l-cos rp signal is therefore'applied to the integrator or pitchcommand computer 308 through a highpass or lead filter to thereby removeany long period or steady state acceleration signals dueto thedifference between the actual and computed normal acceleration. It willbe noted that the normal acceleration or'inertial path damping signal isautomatically insertedY into the autopilot whenever a verticalv pathVmodeY is selected, i.e., when altitude is switched on or when `a glideslope mode is rendered operative. Such switching is accomplished throughswitch K-ll23 (FIG. 1C) closed by energization of vertical path IrelayK-1l2 (FIG. 9).

Furthermore, thedisplacement signal from the altitude control'ltl isalso applied to the input of the pitch command computer 398 which, inthis mode, serves to intef gratethe same to thereby allow any persistentaltitude error to go to Zero. In other words, integration of thealtitude displacement signal through pitch `command computer '368`allows the altitude control mode to be disengaged without a transient.

It will be understood that the operation of the pitch channel'of theautopilot in response to an error signal `from a radio beam, such as anILS glide slope beam,

will be exactly the same as in the altitude control-mode, i.e., that theelevator servo system is controlled in accordance with a glide slopedisplacement signal, the integral of normal acceleration or the rate ofchange in altitude which, in the glide slope, is the sameas the rate ofchange of glide slope error, and the integral of` any long term glideslope error signal.

The automatic pilot of the present inventionm-ay be controlled tolautomatically seek, approach, and thereafter maintain a ilight pathdefined by radio signals. Radio beam guidance facilities fall generallyinto two categories', VOR facilities and lLS facilities. As is known,the former provide en route navigation beams while theA ILS,

of course, provide terminal area or instrument landingV radio guidancebeams, the latter including overlapping.

beams for providing localizer guidance in the rhorizontal plane andglide slope guidance in the vertical plane.

As in the other modes of operation of the autopilot, selection of thedesired mode of operation controls interlock circuits whereby the pilotis conditioned vfor operation in such mode. The following discussion ofthe radio guidance mode will be divided into three sections; VORcoupling, localizer coupling, and finally glide slope coupling. lt willbe noted that in FIGS. 4 and 14 certain of the switching and relaysschematically illustrated in FIG.` 4 have been repeated in FIG. 14, andin the fol*- lowing description reference may be made to either iigurefor those elements common to each.

When the automatic pilot is lirst engaged, the turn selector functionknob i9 is automatically in its Turn Knob position as above described.Under such conditions, D.C. power is applied to lateral beam sensorrelay K-lttl, closing switches K-ltil-l and K--2, the latter preventingenergization of radio beam coupler on relay K-lll. With the selectorknob 19 in the Turn Knob position, the approach coupler is conditionedfor operation upon subsequent turning of the selector knob to theVOR/LOC position.

Assume now that it is desired to approach and main- 1 tain `a VOR beam.Prior to operation of turn selector asegura knob 19, the pilot selectsthe desired omnirange frequency as by frequency selector 403 and setsin, through omniheading selector 404, the bearing of the omniradial hedesires to y. Such omniheading selector is described in detail in U.S.Patent 2,732,550, which is assigned to the same assignee as the presentapplication. As is shown in that patent, the omniheading selector 404provides an output signal which is proportional to the angular deviationof the aircraft from the omniheading. That is, the signal isproportional to the angle between the instantaneous heading of theaircraft and the bearing of the omniradial. Of course, through hisnormal navigation facilities, he may maneuver the craft through the turnknob 14 to a position, as determined by navigation charts, etc., to thevicinity of the selected omnistation. When he desires automatically toapproach `and thereafter maintain the selected radial, he rotates turnfunction selector yknob 19 to the VOR/LOC position. As shown in FIGS. 4and 14, the Selector switch 19 is a make before break type switch sothat such switching will not deenergize the lateral beam sensor relayK-140.

Before describing further the interlock circuitry of FIG. 14, portionsof the beam coupler 400 (FIG. 1A) should be described. Lateral radioreceiver 21,1 provides a signal proportional to the lateral displacementof the craft from the selected radio course. This signal is modulatedand amplied at 405 and applied as an input to a position follow-up servoloop 406 which, through feedback connection from potentiometer 407,positions the followup shaft 408, through amplifier 409, motor 410, andrate generator 411, to a position corresponding to such lateraldisplacement. O11 the output shaft 408 of the approach coupler follow-upservo 406 is the wiper of a sector switch 412. This sector switchcomprises two conducting segments, one of which is relatively long andthe Other fairly short. The approach coup-ler follow-up loop 406 is sodesigned that when the aircraft is located at a distance from the beamgreater than a first predetermined distance (for example, the distanceat which the beam displacement signal is greater than 155 milliamps.),the Wiper of the switch 412 will lie on the large sector appropriatelylabeled plus or minus 155 ,ca Similarly, the small sector ot the switchrepresents a second predetermined lateral distance from the beam (forexample, a distance represented by a displacement signal ofapproximately plus or minus 50` milliamps.) so that when the beam isWithin said predetermined distance from the beam center the wiper willlie on the short sector. This short sector is appropriately labeled plusor minus Silea. The nonconductive portion between the ends of the twosectors therefore represents the distance or displacement of the craftfrom the beam represented by a radio signal having a magniitude between50 pa. and 155 tra.

At this time it will be pointed out that the gain of the radio signalfollow-up loop 406 may be changed under varying beam coupling conditionsas by changing the magnitude of the rate feedback signal from generator411. it should also be mentioned suitable stops on the shaft 40S areprovided sothat the motor cannot drive the sector switch wiper throughan angle greater than 360 in any one direction. Such stops may, forexample, be included within the potentiometer 407, suitable clutch meansbeing provided for preventing damage to the servomotor. Now assume thatit is desired automatically to approach and maintain the radio beamselected and also that the craft is at a distance greater than thatrepresented by a 155 tra. displacement signal. Also assume that thecraft has been placed on a heading which will cause it to intercept thebeam. Switching of function selector switch knob 19 to VOR/LOC positionmerely maintains lateral beam sensor relay K-140 energized throughholding switch l-1401 with the sector switch Wiper, of course, being onthe large sector, as illustrated in FIG. 14. As the craft approaches theedge of the beam, the sector switch arm begins to move towards the endsof the contact sector switch 412. As the craft approaches the beam andthe displacement signal drops to 155 tra., the wiper of switch 412leaves the 155 na. contact sector, thereby deenergizing lateral beamsensor relay K- which, in turn, allows switch K-140-2 to move to itsdeenergize position, thereby energizing radio coupler on relay K-103.With the energization of relay K-103, switch K-103-1 (FIG. 1B) closesand applies the sum of the signal proportional to displacement of thecraft from the beam and a signal proportional to the heading of theaircraft with respect to the beam to roll control channel of theautopilot. The beam displacement signal is taken directly from theoutput of the modulator preamplifier 405 through an amplication stage413 while the heading signal is obtained from the omniheading selector404'. The algebraic sum of these signals is applied to the input ofamplifier limiter 228 in the input to 'the roll command computer 208where the combined signals command a bank angle proportional thereto,just as in the case of a turn rate command signal.

In the limiter circuit, the radio and heading signals are limited so asto limit the bank angle commanded thereby. With the energization ofradio coupler on relay K-103, the coupler may be said to be in a bracketmode, and in this mode the limits imposed by limiter 228 are such thatrelatively large bank angles may be commanded thereby enablingbracketing to occur swiftly. Also, the smooth engage circuit 230 isrendered eifective through smooth engage relay K-119 (FIG. 13) andswitch K- li1t-1 controlled thereby as in the heading select mode, thisbeing accomplished by means of switch K1033. Smooth engage circuit 230is essentially a condenser network which serves to allow the rollcommand signal from the radio coupler to slowly build up over apredetermined time period from Zero to the maximum command value. A timeconstant of about four seconds has been found to be satisfactory forthis buildup. The operation of the automatic pilot in response to thebank angle command is the same as its response to a command either bythe turn rate knob 14 or the heading selector 210 and will not berepeated.

Since the craft is being controlled in accordance with the displacementfrom the beam and the rate of approach as determined by the heading ofthe craft toward the beam, an asymptotic approach path will be executed.As the approach continues, the arm of sector switch 412 will approachthe edge of the short contact sector and when the craft is at a distancerepresented by a signal strength o-f approximately 55 milliamps., thecontact arm will contact this sector. When this occurs, on course relayK-143 will be energized provided that relay switch K-144-1 isdeenergized. The latter switch is controlled from crosscourse velocityrelay K-144 which is energized when the cross-course velocity of theaircraft as determined, in the present VOR coupling mode, by themagnitude of the signal from the omniheading selector which, of course,is a measure of the rate of approach of the craft with respect to thebeam. lf, for some reason, the cross-course velocity exceeds apredetermined value, for example, a velocity resulting from a differencebetween the heading of the craft and the bearing of the beam of, say, 15determined by cross-course velocity sensor 414 which is any suitabiesignal magnitude sensing circuit, such as an amplitier suitably biased,the relay K-lifi will not deenergize and the on course relay willlikewise not be energized, thereby leaving the craft in the bracketcondition. Thus, if the initial heading, air speed, etc., conditionscould not result in an asymptotic approach, the on course mode will notbe instituted and the craft may be allowed to go through an overshootand continue the bracket until, on the next approach, cross-coursevelocity falls below the threshold value as determined by sensor 414.

With K-144 deenergized, on course relay lli-143 becomes energized.Energization of the latter relay will, through switch K-l-S-S (FG. 1B)bypass an attenuator to thereby increase the effective gain of the beamdisplacement signal, and, through switch K-143-2, add the integral ofthe beam error signal through roll integrator 11 whereby to more tightlyhold the craft on the beam and to compensate for the effects of steadycrosswinds tending to blow the craft off course, respectively. Also, oncourse relay K-143 changes the limits imposed on the bank angle commandas by limiter impedance adjustment through switch K-143-6. This changeis in a direction to decrease the bank angle commanded by the radio andheading signals since smaller heading changes are required to maintainthe beam once the craft has acquired the beam. As a further advantage ofsuch bank limiting, a smoother and more comfortable ride is achieved.V

The craft will be maintained on the selected VOR radio through diointeffects of the radio displacement, heading and bank angle signals. Asthe craft approaches the VOR transmitter, the displacement signal willbecome erratic in a region directly over the transmitter, this regionbeing known as the zone of confusion because no sufficiently welldefined radio track information is available for control purposes inthis region. Therefore, it has been found'desirable to sever completelythe control of the aircraft through the radio signal and to leave onlythe heading signal operative to control the craft. This over-the-stationcontrol is disclosed in detail in U.S. Patent No. 2,881,992, which isassigned to the same assignee as the present invention. However, a briefdescription will be included herein for the sake of continuity ofdisclosure.

In FIG. 1A, it will be noted that when on course relay K-143 isenergized (FIG. 14), relay switches K-143-4 and 5 are actuated, theformer serving to connect the output of rate generator 411 to the inputof an over-thestation sensor amplifier 414 and the latter serving toconneet the output of this amplifier to the input of radio beamamplifier 413. As the `displacement term becomes erratic, its ratecomponent as sensed by rate generator 411 will become even more erratic.The over-the-station sensor 414 comprises a circuit which is responsiveto a predetermined magnitude of signal applied to its input and willoperate to supply an output signal of a predetermined magnitude when theinput voltage exceeds its predetermined magnitude. The output of theover-the-station sensor 414 is employed to effectively cut off theoperation of displacement signal amplifier 413 such as by biasing theamplifier -to cut-off or by other means to thereby effectively removethe radio displacement term from the control system during this periodof erratic radio signals. The craft will therefore be caused to maintainthe heading it was on at the time the over-the-station sensor 414 becameeffective. As the craft cornes out of the zone of confusion. the ratecomponent of the displacement signal will drop to a low value, and aftera predetermined time period thereafter, say, 4 seconds, the displacementamplifier 413 will once again be rendered effective to supply the radiodisplacement term to the autopilot for continued control of the craft asit continues outbound on the reciprocal VOR radial under the control ofthe radio displacement, heading, and bank angle signals.

If it is desired to use an ILS approach facility, the turn functionselector knob 19 should be turned to the Turn Knob position, thuspreparing for pilot-inserted maneuvers usual-ly necessary to arrive inthe vicinity of a desired ILS localizer beam. As will become apparent,such operation will also avoid any transients which might otherwiseoccur when tuning to an ILS frequency. When the pilot tunes hisfrequency selector 403 to an ILS frequency, as indicated schematicallyat 415. a pair of ILS relays K126 and K-14S are energized, therebyconditioning the autopilot for an ILS approach. The approach to alocalizer beam is very similar to that to a VOR beam except that in thelocalizer mode the heading is not used as the beam damping or rate ofapproach term, but the actual rate of approach of the craft asdetermined by the rate of change 2@ of the radio displacement signal isused. 'Ihis signal is derived at the output of rate generator 411 in theradio displacement follow-up loop 406.

Energization of ILS relays K-126 and Isl-145 through the selection of anILS frequency serves to energize switches K1261 and K-126-2 whichrespectively supply a beam displacement signal, suitably attenuated foi`localizer bracketing, and a radio rate signal, also suitably attenuatedfor a localizer bracket, which two signals are combined and supplied tolateral beam amplifier 413, the output of which is supplied to theautomatic pilot turn command channel. It will be noted also that ILSrelay K-126-3 serves to remove the heading signal when in the localizermode.

As in the VOR coupling mode, if the craft is considerably displaced fromthe localiZer beam, i.e, in excess of 155 ua., lateral beam sensor relayK-140 will be energized and switch K--2 will keep radio coupler on relayK-103 deenergized. Assuming that the craft is on a heading which willcause it to intercept the beam, when the beam displacement signal dropsto a value below 155 fra., beam sensor relay K-14fl will becomedeenergized, thereby energizing radio coupler on relay K-103 as before.As seen in FIG. 1B, the radio coupler on relay switch K-1031 suppliesthe radio displacement plus rate signal to amplifier limiter 228 to theinput of the roll command computer thereby to cause the craft to roll toan angle determined by the limited radio plus radio rate signal. Theresultant radio displacement, radio rate, and bank angle signals commandan asymptotic approach of the craft to the beam depending upon how farout the approach was initiated. The craft may and very likely will reacha displacement from the beam represented by a 50 ua. radio signal andthe beam sensor switch arm will contact the 50 tra. sector. However, thecraft will continue the approach with no change, i.e., the finalapproach parameters will not be engaged, until the approach relay K-142(FIG. 14) becomes energized. This can occur only under certainconditions. ILS frequency selector relay K-14S fulfills a firstcondition by operation of relay switch K-145-1- The second condition isthat the cross-course velocity must be below a predetermined value, say,2 millivolts per second, thereby deenergizing cross-course velocitydetector Iii-144. The third condition is that the turn knob functionselector must be in its Glide Path position and the glide slope beamintercepted.

ILS frequency relay K-145 also energizes switch K- 145-4 (FIG. 1A) whichsupplies the rate of change of localizer signal to the cross-coursevelocity sensor 414 in place of the heading signal used in the VOR mode.The cross-course velocity sensor is for the purpose of preventingpremature engagement in the approach mode control parameters. Therefore,unless the rate of approach is below said predetermined minimum, theapproach mode cannot be engaged. In'order that the third condition besatisfied, turn function selector knob 19 is rotated to its glide pathposition, thereby arming the coupler for automatic glide path beamcoupling. This aiming is indicated to the pilot by the glide path armlight 416. It will be noted that should the turn function selector knob19 be inadvertently moved to the glide path position when no ILSfunction has in fact been selected, nothing will happen inasmuch as ILSswitch K-145-3 open-circuits the glide slope interlock circuits.

Assuming now that the craft is within the SOua. portion of the beam andis being maintained on the beam by the radio and radio rate signals, andthat the craft is approaching the glide slope beam from the undersidethereof. Vertical beam sensor 414 (this may be the same circuit employedfor the over-the-station sensor) will prevent premature engagement ofthe glide slope control and also premature switching of the lateralchannel to its approach parameters until the aircraft is about tointercept the center of the glide slope beam. The vertical beam sensorderives its input signal from the glide path receiver 311 through nowclosed relay switch K-143-4. Its output, when the glide slope beam errorsignal drops below a predetermined value, for example, 15 millivolts,controls vertical beam sensor relay K-146 which, in turn, energizesglide path engage relay K-147. The latter relay fulfills the thirdcondition mentioned above by closing switch K-147-2 in the approachrelay K-142 energization circuit. As approach relay K-142* is energized,a holding circuit, through ILSvrelay switch VK- -iS-S and approach relayswitch K-142`-1, is established so that further operation of thecross-course velocity detector switch K-144-1 will not deenergize theapproach relay. It will also be noted that a second approach relay K-122is energized through approach relay switch K-142-2.

The craft is now in its final approach maneuver and in order to providevery tight beam coupling various gains in the approach coupler arechanged. In the first place, the gain of the displacement follow-up loop406 is changed by operation of approach relay switch K-12Z-1 in adirection to increase the response of the follow-up loop to lateral beamdisplacements. At the same time, the magnitude of the displacementsignal supplied to the autopilot through amplifier 413` is increasedthrough approach relay switch K-122-3, and finally, the magnitude of therate signal supplied to the autopilot is also increased throughoperation of approach relay switch K- 122-2. ln this manner, the craftis controlled to precisely follow the localizer beam during the approachmode. Since large bank angles are to be avoided in the final approachconiiguration, the bank limiter is adjusted by means of switch K-122-4to decrease the limits imposed upon the sum of the radio and radio ratesignals, that is, to reduce the magnitude of the sum signal which it canpass. it will be further noted that any standoff errors which mayaccumulate during the approach are eliminated through the operation ofthe roll integrator 11 which is rendered responsive to any standoifdisplacement error through approach relay switch K-142-3.

in order to increase the yaw stability of the craft during the approachmaneuver, a signal from the directional gyro 107 (FIG. 1B) is insertedintoy the rudder channel in such a manner that short term yaw gusts areopposed while long term turn commands by the radio are unopposed. Thisis accomplished through yaw followup high tachometer gradient relayK-171 (FIG. 14) which is energized simultaneously with the energizationof approach relay K-122. Energization of this relay serves throughswitch K`171-1 to decrease the gain through the yaw follow-up loop 108so that it can readily follow up on long term gyro signals, yet at thesame time 'oe unable to follow up on short term gyro signals. Thus, anyshort term deviations of the craft in yaw, such as produced by lateralgusts7 etc., are fed directly to the rudder servo system throughapproach relay switch Krim-4.

During the approach mode, the elevator channel of the autopilot iscontrolled in accordance with displacement of the craft vertically fromthe center of the glide slope beam. Any deviation of the craft from thebeam is detected by glide slope receiver 311 which supplies an errorsignal to the elevator servo system through amplifier limiter' 321 whichoperates in the glide slope mode, exactly as in the altitude controlmode, to provide a displacement reference command signal for theelevator servo system. Additionally, and again as in the altitudecontrol mode, any persistent. glide slope displacement signal isintegrated out through the operation of the pitch command computer 392which has been placed in its integrator mode of operation. The latter isaccomplished through vertical path relay Y-112 becoming energized whenglide path engage relay K-147 became energized, i.e., through switchK-147-4 (FIG. 9); this, in turn, energizing pitch computer gear shiftsolenoid MG-101. inertial path damping is also provided in the approachmode as in the altitude control mode through the operation of verticalpath relay lil-i12 through its switch 14412-3, which supplies a signalin accordance with the vertical acceleration of the craft, the lattersignal being supplied to the pitch command computer 302 where it isintegrated and appears in its output as a rate of change of heightsignal. Therefore, the craft is controlled to precisely follow the glideslope beam through the combined effect of the displacement of the crafton the beam, the rate of change of height of the craft, and the timeintegral of the displacement.

It will be noted that in the radio coupling modes of operation of theautopilot, the switching sequences which occur automatically frominitial bracketing to on course or approach are dependent solely uponthe position and/ or movement of the craft with respect to the beam,thereby providing a very positive coupling to the beam.

From the foregoing specification, it will be clear that the autopilotdisclosed herein is similar in a great many respects to that disclosedin the above-identified copending parent application and yet there areherein disclosed many novel features not specifically set forth in theparent application. However, although specific and detailed disclosuresof the autopilot of the present invention have been set forth, it isclearly apparent that many changes could be made in this specificconstruction and many widely dierent embodiments could be constructedWithout departing from the true scope and spirit thereof. Therefore, itis intended that all matter contained in the above description or shownin the accompanying drawings should be interpreted as illustrative andnot in a limiting sense.

What is claimed is:

l. In a yaw damper for an aircraft having a rudder and servomotor meansfor controlling the rudder whereby to control yawing of the aircraft,the combination comprising first accelerometer means for supplying afirst signal in accordance with the acceleration of the aircraft aboutits yaw axis and a second signal in accordance with the lateralacceleration of said aircraft, second accelerometer means for supplyinga third signal in accordance with the angular acceleration of the craftabout its roll axis, integrator means responsive to said second andthird signals for supplying a fourth signal in accordance with the timeintegral of the sum thereof, and means responsive to said first, secondand fourth signals for controlling said servomotor means in accordancewith the sum thereof.

2. A yaw damper for an aircraft having a rudder for controlling yawingthereof, comprising a servomotor for controlling the operation of saidrudder, first accelerometer means for supplying rst and second signalsin accordance with yaw and lateral accelerations of the aircraftrespectively, second accelerometer means for supplying a signal inaccordance with roll accelerations of the aircraft, means for combiningsaid lateral and roll acceleration signals for supplying an output inaccordance with the sum thereof, integrator means responsive to saidoutput for supplying a signal in accordance with the time integralthereof, means for combining said yaw acceleration and integral signals,means responsive to the resultant -thereof and to said lateralacceleration signal for supplying a servomotor control signal, and meansfor supplying said control signal to said servomotor means.

3. Apparatus as set forth in claim 2 wherein said rst accelerometermeans comprises two accelerometers, one located forward of the craftscenter of gravity and the other located substantially at the craftscenter of gravity, for supplying signals in accordance with craftaccelerations at these locations, and means for combining the signalsfrom both said accelerometers for supplying said yaw accelerationsignal.

4. A yaw damper for an aircraft having a yaw control surface, comprisingmeans for supplying a signal pro portional to angular acceleration ofsaid craft about the provide both short and long term coordination ofmanual-A ly commanded turns during yaw damper operation.

References Cited in the le of this patent UNITED STATES PATENTS CheneryNov. 8, Alderson Mar. 18, Owen Apr. 8, 7 Kellogg Nov. 11, GilleA Dec. 9,Kerpchar Apr. 21, Noxon July 7,

